This invention relates generally to methods of making gas turbine components and more particularly relates to methods of making improved combustors and/or transition pieces for utilization in gas turbines where film cooling may be extremely limited or might not even be possible.
Traditional gas turbine combustors use diffusion (i.e., nonpremixed) flames in which fuel and air enter the combustion chamber separately. The process of mixing and burning produces flame temperatures exceeding 3900 degrees F. Since conventional combustors and/or transition pieces having liners are generally capable of withstanding for about ten thousand hours (10,000) a maximum temperature on the order of only about 1500 degrees F., steps to protect the combustor and/or transition piece liners must be taken. This has typically been done by film-cooling which involves introducing the relatively cool compressor air into a plenum surrounding the outside of the combustor. In this prior arrangement, the air from the plenum passes through louvers in the combustor liner and then passes as a film over the inner surface of the combustor liner, thereby maintaining combustor liner integrity.
Because diatomic nitrogen rapidly disassociates at temperatures exceeding about 3000.degree. F. (about 1650.degree. C.), the high temperatures of diffusion combustion result in relatively large NO.sub.x emissions. One approach to reducing NOx emissions has been to premix the maximum possible amount of compressor air with fuel. The resulting lean premixed combustion produces cooler flame temperatures and thus lower NOx emissions. Although lean premixed combustion is cooler than diffusion combustion, the flame temperature is still too hot for prior conventional combustor liners to withstand.
Furthermore, because the advanced combustors premix the maximum possible amount of air with the fuel for NOx reduction, little or no cooling air is available making film-cooling of the combustor liner impossible. Thus, means such as thermal barrier coating in conjunction with "backside" cooling have been considered to protect the combustor liner from destruction by such high heat. Backside cooling involved passing the compressor air over the outer surface of the combustor liner prior to premixing the air with the fuel.
Lean premixed combustion reduces NO.sub.x emissions by producing lower flame temperatures. However, the lower temperatures, particularly along the inner surface or wall of the combustor liner, tend to quench oxidation of carbon monoxide and unburned hydrocarbons and lead to unacceptable emissions of these species. To oxidize carbon monoxide and unburned hydrocarbons, a liner would require a thermal barrier coating of extreme thickness (50-100 mils) so that the surface temperature could be high enough to ensure complete burnout of carbon monoxide and unburned hydrocarbons. This would be approximately 1800-2000 degrees F. bond coat temperature and approximately 2200 degrees F. TBC (Thermal Barrier Coating) temperature for combustors of typical lengths and flow conditions. However, such thermal barrier coating thicknesses and temperatures for typical gas turbine component lifetimes are beyond current materials known capabilities. Known thermal barrier coatings degrade in unacceptably short times at these temperatures and such thick coatings are susceptible to spallation.
Advanced cooling concepts now under development require the fabrication of complicated cooling channels in thin-walled structures. The more complex these structures are, the more difficult they are to make using conventional techniques, such as casting. Because these structures have complexity and wall dimensions that may be beyond the castability range of advanced superalloys, and which may exceed the capabilities of the fragile ceramic cores used in casting, both in terms of breakage and distortion, new methods of fabricating must be developed to overcome these prior limitations. Possible geometries for enhanced cooling are disclosed in Docket No. 51DV5608PA, the disclosure of which is incorporated herein by reference.
Because the accuracy of placement of cooling features and wall thicknesses is much greater than for ceramic-cored castings, the powder foil process, described in U.S. Pat. Nos. 5,427,736 and 5,480,468, the disclosure of each is herein incorporated by reference, may provide for greatly reduced tolerances on wall thicknesses. These capabilities are also of interest in fabricating components other than airfoils, such as combustors, which can be cylindrical, and more irregularly shaped components such as transition pieces. In those kinds of structures, wrought alloys may be used, with considerable brazing and/or welding of cooling features (such as impingement sleeves) being required. Accuracy of placement of such cooling features, and retention of mechanical behavior of strong superalloys and avoidance of structural damage during welding, may be a severe limitation to materials that can be considered.
More efficient cooling structures may allow significant cooling flow reduction, perhaps up to 60%, without increasing the metal temperature. These reductions may also be realized with a combination of new cooling structures and other system/cycle changes.
Accordingly, there is a need for new and improved methods of making gas turbine components, such as, for example, combustors/transition pieces which can withstand combustion temperatures without film-cooling and yet maintain flame stability and burn out carbon monoxide and unburned hydrocarbons, such as advanced cooling concepts for low-emissions-combustors and transition pieces (particularly NO.sub.x emissions). Such methods should produce combustor/transition pieces having cooling channels in thin-walled structures which allow the inner surface of the combustor/transition piece to maintain reasonable metal temperatures. Efficient cooling combustor/transition piece structures should provide cooling flow reduction of about thirty-five percent (35%) to about sixty percent (60%) without increasing the metal temperature of the inner surface of the combustor and should have internal features such as turbulation promoters which must have sharper internal edges than can be currently produced by casting. Because the accuracy of placement of cooling features and wall thicknesses can be much greater than for ceramic-cored castings, utilizing a powder foil process should provide for greatly reduced tolerances on wall thicknesses, and should also provide for the production of sharp-edged internal features. These capabilities could also be used in fabricating components, such as combustors, which can be cylindrical, and more irregularly shaped components such as transition pieces. In those kinds of structures, wrought alloys might be used, with considerable brazing and/or welding of cooling features (such as impingement sleeves) being required. Accuracy of placement of such cooling features, and retention of mechanical behavior of strong superalloys and avoidance of structural damage during welding, may be a severe limitation to materials that can be considered.